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NASA NTRS2020

Hardware Verification and Validation for a Navigation Sensor Software Model in Support of Flight Vehicle Performance Analysis

… or, “It’s in the details, how to make complicated software perform like complicated hardware.” In attempts to minimize development time and quickly build an operational vehicle, NASA’s Space Launch System (SLS) has had to be intentional about integrated testing. Constraints on budget and schedule have required balance between testing needs and the desire for an integrated flight vehicle as soon as possible. To provide key insights early in design and analysis cycles, a large amount of effort has shifted into maturing and validating models at the component level with integrated testing as a means to validate their integration. In terms of SLS Navigation, this, and the model-based design approach have pushed explicit requirements for sensor models to be validated against flight hardware to high precision. This paper covers the approach taken to verify and validate the models for the two key navigation sensors on the SLS vehicle, the Redundant Inertial Navigation Sensor and the Rate Gyro Assembly. These models are used in performance evaluation, fault detection, and operations development extensively. Using a mix of data from hardware vendor documentation and testing reports, limited in-house testing, and integration activities, these models were able to be validated against flight hardware at multiple levels, from the internal software design to statistical behavior at the raw sensor and integrated box levels. The high level of insight into the hardware elements is instrumental to support flight certification activities and building confidence in SLS Navigation capability. Focused testing enabled additional insight and validation that proved invaluable and the resulting insights were used to focus and mature models. Additionally, of having validated performance-based hardware models enables a wide breadth of activities including detailed fault detection studies and integration into future vehicle frameworks, such as an upper stage and provide a valuable asset to continued SLS analysis and design.

Evan J Anzalone↗

NASA NTRS2020

Hardware Verification and Validation for a Navigation Sensor Software Model in Support of Flight Vehicle Performance Analysis

Thomas Park↗

NASA NTRS1990

Entry vehicle performance analysis and atmospheric guidance algorithm for precision landing on Mars

Future missions to Mars may require pin-point landing precision, possibly on the order of tens of meters. The ability to reach a target while meeting a dynamic pressure constraint to ensure safe parachute deployment is complicated at Mars by low atmospheric density, high atmospheric uncertainty, and the desire to employ only bank angle control. The vehicle aerodynamic performance requirements and guidance necessary for 0.5 to 1.5 lift drag ratio vehicle to maximize the achievable footprint while meeting the constraints are examined. A parametric study of the various factors related to entry vehicle performance in the Mars environment is undertaken to develop general vehicle aerodynamic design requirements. The combination of low lift drag ratio and low atmospheric density at Mars result in a large phugoid motion involving the dynamic pressure which complicates trajectory control. Vehicle ballistic coefficient is demonstrated to be the predominant characteristic affecting final dynamic pressure. Additionally, a speed brake is shown to be ineffective at reducing the final dynamic pressure. An adaptive precision entry atmospheric guidance scheme is presented. The guidance uses a numeric predictor-corrector algorithm to control downrange, an azimuth controller to govern crossrange, and analytic control law to reduce the final dynamic pressure. Guidance performance is tested against a variety of dispersions, and the results from selected tests are presented. Precision entry using bank angle control only is demonstrated to be feasible at Mars.

Dieriam, Todd A.↗

NASA NTRS2006

The Vehicle Integrated Performance Analysis Experience: Reconnecting With Technical Integration

Very early in the Space Launch Initiative program, a small team of engineers at MSFC proposed a process for performing system-level assessments of a launch vehicle. Aimed primarily at providing insight and making NASA a smart buyer, the Vehicle Integrated Performance Analysis (VIPA) team was created. The difference between the VIPA effort and previous integration attempts is that VIPA a process using experienced people from various disciplines, which focuses them on a technically integrated assessment. The foundations of VIPA s process are described. The VIPA team also recognized the need to target early detailed analysis toward identifying significant systems issues. This process is driven by the T-model for technical integration. VIPA s approach to performing system-level technical integration is discussed in detail. The VIPA process significantly enhances the development and monitoring of realizable project requirements. VIPA s assessment validates the concept s stated performance, identifies significant issues either with the concept or the requirements, and then reintegrates these issues to determine impacts. This process is discussed along with a description of how it may be integrated into a program s insight and review process. The VIPA process has gained favor with both engineering and project organizations for being responsive and insightful

McGhee, D. S.↗

NASA NTRS2005

Vehicle Integrated Performance Analysis, the VIPA Experience: Reconnecting with Technical Integration

Today's NASA is facing significant challenges and changes. The Exploration initiative indicates a large increase in projects with limited increase in budget. The Columbia report has criticized NASA for its lack of insight and technical integration impacting its ability to provide safety. The Aldridge report is advocating NASA find new ways of doing business. Very early in the Space Launch Initiative (SLI) program a small team of engineers at MSFC were asked to propose a process for performing a system level assessment of a launch vehicle. The request was aimed primarily at providing insight and making NASA a "smart buyer." Out of this effort the VIPA team was created. The difference between the VIPA effort and many integration attempts is that VIPA focuses on using experienced people from various disciplines and a process which focuses them on a technically integrated assessment. Most previous attempts have focused on developing an all encompassing software tool. In addition, VIPA anchored its process formulation in the experience of its members and in early developmental Space Shuttle experience. The primary reference for this is NASA-TP-2001-210092, "Launch Vehicle Design Process: Characterization, Technical Integration, and Lessons Learned," and discussions with its authors. The foundations of VIPA's process are described. The VIPA team also recognized the need to drive detailed analysis earlier in the design process. Analyses and techniques typically done in later design phases, are brought forward using improved computing technology. The intent is to allow the identification of significant sensitivities, trades, and design issues much earlier in the program. This process is driven by the T-model for Technical Integration described in the aforementioned reference. VIPA's approach to performing system level technical integration is discussed in detail. Proposed definitions are offered to clarify this discussion and the general systems integration dialog. VIPA's capabilities and process can now be used to significantly enhance the development and monitoring of realizable project requirements. This is done through the use of VIPA's V-model. Starting with a given concept, VIPA's assessment validates the concept's stated performance, identifies significant issues either with the concept or the requirements, and then re-integrates these issues to determine impacts. This process is discussed along with a description of how it may be integrated into a program's insight and review process. The VIPA process has been employed successfully on several projects including SLI, Orbital Space Plane (OSP), and several heavy lift concepts for Exploration. It has also been proposed for use on the Jupiter Icy Moon (JIMO) spacecraft. The VIPA process has gained favor with both engineering and project organizations for being responsive and insightful. Selected results from these assessments will be presented.

McGhee, David S.↗

NASA NTRS1988

Mars Exploration: Mission and System Analysis of the Split Sprint Mission

The information presented in this report provides detailed mission and system analysis of the Split Sprint Mission to Mars. These data were generated from September 1987 through July 1988 and involved numerous individuals within Program Development. This mission derives its name from the fact that the total mission payload is split between two vehicles, a cargo vehicle, and a piloted vehicle. The cargo vehicle carries the Mars lander and propellant for the piloted vehicle return to Earth. Then approximately 1 1/2 years later, the piloted vehicle carries the crew on a higher energy "sprint” trajectory to Mars. The advantages of the Split Sprint Mission is that the payload for the piloted vehicle is minimized and the crew would be in space for only 14 months. Mission and system data are presented for a Sprint Mission in the year 2004. Many important aspects of the mission were studied in detail. These include orbital lighting and lifetime analysis, Mars parking orbit selection, Earth departure window analysis, lander descent and ascent performance, vehicle aerobraking requirements, mission orbital maneuvers, and many other analyses listed in the Table of Contents. Emphasis was placed on vehicle performance analysis which ranged from recovery of the Earth departure stages to aerocapture of the Earth Return Capsule. There is also a section describing the computer programs that were used for this study. These programs provide many useful tools for future studies of manned planetary missions. This introductory section will give an overall view of the Split Sprint Mission and describe the orbital operations and vehicles that were studied.

Mars Architecture Team↗

NASA NTRS1982

Transistorized PWM inverter-induction motor drive system

This paper describes the development of a transistorized PWM inverter-induction motor traction drive system. A vehicle performance analysis was performed to establish the vehicle tractive effort-speed requirements. These requirements were then converted into a set of inverter and motor specifications. The inverter was a transistorized three-phase bridge using General Electric power Darlington transistors. The description of the design and development of this inverter is the principal object of this paper. The high-speed induction motor is a design which is optimized for use with an inverter power source. The primary feedback control is a torque angle control with voltage and torque outer loop controls. A current-controlled PWM technique is used to control the motor voltage. The drive has a constant torque output with PWM operation to base motor speed and a constant horsepower output with square wave operation to maximum speed. The drive system was dynamometer tested and the results are presented.

Peak, S. C.↗

NASA NTRS1982

Improved transistorized AC motor controller for battery powered urban electric passenger vehicles

An ac motor controller for an induction motor electric vehicle drive system was designed, fabricated, tested, evaluated, and cost analyzed. A vehicle performance analysis was done to establish the vehicle tractive effort-speed requirements. These requirements were then converted into a set of ac motor and ac controller requirements. The power inverter is a three-phase bridge using power Darlington transistors. The induction motor was optimized for use with an inverter power source. The drive system has a constant torque output to base motor speed and a constant horsepower output to maximum speed. A gear shifting transmission is not required. The ac controller was scaled from the base 20 hp (41 hp peak) at 108 volts dec to an expanded horsepower and battery voltage range. Motor reversal was accomplished by electronic reversal of the inverter phase sequence. The ac controller can also be used as a boost chopper battery charger. The drive system was tested on a dynamometer and results are presented. The current-controlled pulse width modulation control scheme yielded improved motor current waveforms. The ac controller favors a higher system voltage.

Peak, S. C.↗

NASA NTRS2000

GPS Sounding Rocket Development at NASA with Simultaneous Multi-Payload Tracking Application

An inverse differential GPS system has been developed for Sounding Rocket use which includes the flight unit and a ground station capable of extracting GPS data from sounding rocket telemetry, performing a real time differential solution and graphically displaying the rocket's path relative to a predicted trajectory plot. Accuracy has been proven to within less than 10 meters. Postprocessing has increased the precision to within 10 - 20 centimeters. The system has been successfully flown several times and delivered to the Sounding Program Office for routine field use. In addition to providing position, velocity and time GPS data has been used on sounding rockets for vehicle performance analysis, effecting a one hundred fold improvement in data time tagging, and steering an optical tracking device to intercept payloads launched from over the horizon. Precise velocity separation information and timing has been provided to multiple payload systems. Future plans include its use for Range Safety and enabling of interferometric techniques. The technology and software developed also has potential application to small satellite navigation and formation flying.

Bull, Barton↗

NASA NTRS2011

Orbiter Trajectory Analysis for a Two-Stage Reusable Launch Vehicle

Trajectory analysis performed on NASA's reference two-stage-to-orbit launch vehicle upper stage will be presented. The work was completed in support of the Hypersonics Multidisciplinary Analysis and Optimization effort for the NASA-Air Force Joint System Study. Three degree-of-freedom (3-DOF) untrimmed trajectory analysis was performed for the orbiter ascent, closure and re-entry. An iterative closure process resulted in a 333,000 lb initial mass for the orbiter. The re-entry trajectory satisfied heating constraints for all payload out cases and met the constraints with reduced margins for payload in cases. Abort trajectories for engine out at staging, engine out during ascent, and failure to circularize in orbit, gave insight to the robustness of the orbiter. A trimmed ascent trajectory defined an engine gimbal location and the body flap angle best suited for maximizing injected mass. A trimmed re-entry trajectory revealed a need to update the trim routine to accommodate full flap aerodynamic data.

Cowling, Adam L.↗

NASA NTRS1989

Wave combustors for trans-atmospheric vehicles

A performance analysis is given of a conceptual transatmospheric vehicle (TAV). The TAV is powered by a an oblique detonation wave engine (ODWE). The ODWE is an airbreathing hypersonic propulsion system which utilizes shock and detonation waves to enhance fuel-air mixing and combustion in supersonic flow. In this wave combustor concept, an oblique shock wave in the combustor can act as a flameholder by increasing the pressure and temperature of the air-fuel mixture, thereby decreasing the ignition delay. If the oblique shock is sufficiently strong, then the combustion front and the shock wave can couple into a detonation wave. In this case, combustion occurs almost instantaneously in a thin zone behind the wave front. The result is a shorter lighter engine compared to the scramjet. The ODWE-powered hypersonic vehicle performance is compared to that of a scramjet-powered vehicle. Among the results outlined, it is found that the ODWE trades a better engine performance above Mach 15 for a lower performance below Mach 15. The overall higher performance of the ODWE results in a 51,000-lb weight savings and a higher payload weight fraction of approximately 12 percent.

Menees, Gene P.↗

NASA NTRS1990

Vehicle processing operations database

The Vehicle Processing Operations Data base (VPOD) has been designed and constructed to enable the collection and preliminary analysis of data relative to the processing of planetary exploration vehicles in low Earth orbit. VPOD has the capability to hierarchically store the events, assemblies, and equipment data needed for performing vehicle processing analysis, to whatever level of detail available. This permits some analysis to be performed even during the very early phases of vehicle and mission design. As more data about the events, assemblies, or equipment become available, the analysis can then be refined using the increased level of detail. VPOD output is based on a mission defined by selecting from the VPOD Events data base the events that are required to perform the desired tasks. The outputs available include total crew time, crew skill requirements, equipment needed, assemblies worked per event, to any available level of detail. The VPOD design is structured to interface with an operations simulation capability which is currently in the requirements definition phase. The simulator will use information from the VPOD to perform operations analysis for missions being conducted in low Earth orbit. VPOD information, along with information from the science missions operations data base and a Space Station configuration data base, will be used by the simulator to provide a capability to study total system operations and resources requirements.

Ganoe, George G.↗

NASA NTRS2007

Crew Exploration Vehicle Launch Abort Controller Performance Analysis

This paper covers the simulation and evaluation of a controller design for the Crew Module (CM) Launch Abort System (LAS), to measure its ability to meet the abort performance requirements. The controller used in this study is a hybrid design, including features developed by the Government and the Contractor. Testing is done using two separate 6-degree-of-freedom (DOF) computer simulation implementations of the LAS/CM throughout the ascent trajectory: 1) executing a series of abort simulations along a nominal trajectory for the nominal LAS/CM system; and 2) using a series of Monte Carlo runs with perturbed initial flight conditions and perturbed system parameters. The performance of the controller is evaluated against a set of criteria, which is based upon the current functional requirements of the LAS. Preliminary analysis indicates that the performance of the present controller meets (with the exception of a few cases) the evaluation criteria mentioned above.

Sparks, Dean W., Jr.↗

NASA NTRS1976

Performance analysis of a laser propelled interorbital tansfer vehicle

Performance capabilities of a laser-propelled interorbital transfer vehicle receiving propulsive power from one ground-based transmitter was investigated. The laser transmits propulsive energy to the vehicle during successive station fly-overs. By applying a series of these propulsive maneuvers, large payloads can be economically transferred between low earth orbits and synchronous orbits. Operations involving the injection of large payloads onto escape trajectories are also studied. The duration of each successive engine burn must be carefully timed so that the vehicle reappears over the laser station to receive additional propulsive power within the shortest possible time. The analytical solution for determining these time intervals is presented, as is a solution to the problem of determining maximum injection payloads. Parameteric computer analysis based on these optimization studies is presented. The results show that relatively low beam powers, on the order of 50 MW to 60 MW, produce significant performance capabilities.

Minovitch, M. A.↗

NASA NTRS2005

Integration of Launch Vehicle Simulation/Analysis Tools and Lunar Cargo Lander Design

Simulation and analysis of vehicle performance is essential for design of a new launch vehicle system. It is more and more demand to have an integrated, highly efficient, robust simulation tool with graphical user interface (GUI) for vehicle performance and simulations. The objectives of this project are to integrate and develop launch vehicle simulation and analysis tools in MATLAB/Simulink under PC Platform, to develop a vehicle capable of being launched on a Delta-IV Heavy Launch Vehicle which can land on the moon with the goal of pre-implanting cargo for a new lunar mission, also with the capability of selecting other launch vehicles that are capable of inserting a payload into Trans-Lunar Injection (TLI). The vehicle flight simulation software, MAVERIC-II (Marshall Aerospace VEhicle Representation In 'C'), developed by Marshall Space Flight Center was selected as a starting point for integration of simulation/analysis tools. The goals are to convert MAVERIC-II from UNIX to PC platform and build input/output GUI s in the MATLAB environment, and then integrate them under MATLAB/Simulink with other modules. Currently, MAVERIC-II has been successfully converted from UNIX to PC using Microsoft Services for UNIX subsystem on PC. Input/Output GUI's have been done for some key input/output files. Calling MAVERIC-II from Simulink has been tested. Details regarding Lunar Cargo Lander Design are described in Part 2/2 of the paper on page X-1.

Shiue, Yeu-Sheng Paul↗

NASA NTRS2015

DUKSUP: A Computer Program for High Thrust Launch Vehicle Trajectory Design and Optimization

From the late 1960s through 1997, the leadership of NASAs Intermediate and Large class unmanned expendable launch vehicle projects resided at the NASA Lewis (now Glenn) Research Center (LeRC). One of LeRCs primary responsibilities --- trajectory design and performance analysis --- was accomplished by an internally-developed analytic three dimensional computer program called DUKSUP. Because of its Calculus of Variations-based optimization routine, this code was generally more capable of finding optimal solutions than its contemporaries. A derivation of optimal control using the Calculus of Variations is summarized including transversality, intermediate, and final conditions. The two point boundary value problem is explained. A brief summary of the codes operation is provided, including iteration via the Newton-Raphson scheme and integration of variational and motion equations via a 4th order Runge-Kutta scheme. Main subroutines are discussed. The history of the LeRC trajectory design efforts in the early 1960s is explained within the context of supporting the Centaur upper stage program. How the code was constructed based on the operation of the AtlasCentaur launch vehicle, the limits of the computers of that era, the limits of the computer programming languages, and the missions it supported are discussed. The vehicles DUKSUP supported (AtlasCentaur, TitanCentaur, and ShuttleCentaur) are briefly described. The types of missions, including Earth orbital and interplanetary, are described. The roles of flight constraints and their impact on launch operations are detailed (such as jettisoning hardware on heating, Range Safety, ground station tracking, and elliptical parking orbits). The computer main frames on which the code was hosted are described. The applications of the code are detailed, including independent check of contractor analysis, benchmarking, leading edge analysis, and vehicle performance improvement assessments. Several of DUKSUPs many major impacts on launches are discussed including Intelsat, Voyager, Pioneer Venus, HEAO, Galileo, and Cassini.

Launch vehicle performance optimization↗

NASA NTRS2014

DUKSUP: A Computer Program for High Thrust Launch Vehicle Trajectory Design and Optimization

From the late 1960's through 1997, the leadership of NASA's Intermediate and Large class unmanned expendable launch vehicle projects resided at the NASA Lewis (now Glenn) Research Center (LeRC). One of LeRC's primary responsibilities --- trajectory design and performance analysis --- was accomplished by an internally-developed analytic three dimensional computer program called DUKSUP. Because of its Calculus of Variations-based optimization routine, this code was generally more capable of finding optimal solutions than its contemporaries. A derivation of optimal control using the Calculus of Variations is summarized including transversality, intermediate, and final conditions. The two point boundary value problem is explained. A brief summary of the code's operation is provided, including iteration via the Newton-Raphson scheme and integration of variational and motion equations via a 4th order Runge-Kutta scheme. Main subroutines are discussed. The history of the LeRC trajectory design efforts in the early 1960's is explained within the context of supporting the Centaur upper stage program. How the code was constructed based on the operation of the Atlas/Centaur launch vehicle, the limits of the computers of that era, the limits of the computer programming languages, and the missions it supported are discussed. The vehicles DUKSUP supported (Atlas/Centaur, Titan/Centaur, and Shuttle/Centaur) are briefly described. The types of missions, including Earth orbital and interplanetary, are described. The roles of flight constraints and their impact on launch operations are detailed (such as jettisoning hardware on heating, Range Safety, ground station tracking, and elliptical parking orbits). The computer main frames on which the code was hosted are described. The applications of the code are detailed, including independent check of contractor analysis, benchmarking, leading edge analysis, and vehicle performance improvement assessments. Several of DUKSUP's many major impacts on launches are discussed including Intelsat, Voyager, Pioneer Venus, HEAO, Galileo, and Cassini.

high thrust trajectory design↗

NASA NTRS1995

Advanced Transportation System Studies Technical Area 2 (TA-2) Heavy Lift Launch Vehicle Development Contract: Technical Results - Volume 2

The purpose of the Advanced Transportation System Studies (ATSS) Technical Area 2 (TA-2) Heavy Lift Launch Vehicle Development contract was to provide advanced launch vehicle concept definition and analysis to assist NASA in the identification of future launch vehicle requirements. Contracted analysis activities included vehicle sizing and performance analysis, subsystem concept definition, propulsion subsystem definition (foreign and domestic), ground operations and facilities analysis, and life cycle cost estimation. This document is Volume 2 of the final report for the contract. It provides documentation of selected technical results from various TA-2 analysis activities, including a detailed narrative description of the SSTO concept assessment results, a user's guide for the associated SSTO sizing tools, an SSTO turnaround assessment report, an executive summary of the ground operations assessments performed during the first year of the contract, a configuration-independent vehicle health management system requirements report, a copy of all major TA-2 contract presentations, a copy of the FLO launch vehicle final report, and references to Pratt & Whitney's TA-2 sponsored final reports regarding the identification of Russian main propulsion technologies.

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